This invention relates generally to gas turbine engines and more particularly, to methods and apparatus for cooling gas turbine engine rotor assemblies.
A typical gas turbine engine includes a rotor assembly having circumferentially-spaced rotor blades. Each rotor blade, sometimes referred to as a bucket, includes an airfoil that extends radially outward from a platform. Each rotor blade also includes a dovetail that extends radially inward from a shank extending between the platform and the dovetail. The dovetail is used to mount the rotor blade within the rotor assembly to a rotor disk or spool. Known blades are hollow such that an internal cooling cavity is defined at least partially by the airfoil, platform, shank, and dovetail.
With respect to gas turbine operation, increasing inlet firing temperatures provides improved output and engine efficiencies. Increasing the inlet firing temperature results in increased gas path temperatures. Such increased gas path temperatures can result in added stress to the bucket platforms, including possibly oxidation, creep and cracking. Further, in gas turbines where closed loop cooling circuits are used in upstream airfoil components, there is no film cooling and therefore the downstream bucket platforms do not have the benefit from the film carryover from the upstream airfoils. This exacerbates the potential distress on the bucket platforms.
Some recent known turbine blade configurations do utilize film cooling for cooling the blade platform. Specifically, compressor discharge air is routed through an opening or openings in the platform, and a layer of cooling film forms on the platform to protect the platform from the high flow path temperatures. With such film cooling, however, there may only be sufficient pressure to film cool the aft section of the platform where the flow path air has been accelerated to drop the local static pressure.